The surfaces of turbine engine parts are exposed to the hot gases from the turbine combustion process. Turbine engine superalloy materials are selected based on their high temperature stability and corrosion resistance. Well-known superalloys, for example nickel based superalloys such as Inconel™ 718, Inconel™ 722 and Udimet™ 720 demonstrate good resistance to oxidation and corrosion damage. However even these materials experience degradation under severe conditions at high temperatures. Oxidation and corrosion reactions at the surface of the component parts can cause metal wastage and loss of wall thickness. The loss of metal rapidly increases the stresses on the respective component part and can ultimately result in part failure. Protective overlays are thus applied to these component parts to protect them from degradation by oxidation and corrosion.
Various corrosion-resistant layers and multilayer overlay systems have been suggested and used to protect turbine engine components, particularly compressor rotor blades. Assessment of the prior art overlay systems have revealed general deficiencies in their functional properties and appearance, as well as several possible failure modes.
For example, a prior art commercially available multilayer overlay system is designed for lower service temperatures and provides effective protection up to 1200° F. However this prior art overlay system would be prone to cracking and delamination at elevated operating temperatures (≧˜1300° F.) of newer engines if it were used on such advanced engines. FIG. 1 shows delamination of the prior art overlay system from Inconel™ 718 substrate exposed to 1400° F. for 145 hrs, which is at a temperature significantly above its designed operating temperatures.
FIGS. 2A and 2B illustrate[s] other issues or problems associated with prior art multilayer overlay systems. The prior art coated substrates in FIGS. 2A and 2B show a “gritty” coating appearance (i.e. visible particle inclusions). These particle inclusions were observed after application of intermediate layers and tend to become more pronounced after application of the seal coat layer. These defects were attributed to external contamination during layer application, such as airborne contaminants, surface irregularities, etc.
Other type of possible issues or problems that may be associated with the prior art based overlay systems are the 1 mm to 3 mm diameter round spots (i.e. “white spots”) on some parts coated with the prior art overlay system. As seen in FIGS. 2A and 2B, the “white spots” appear much lighter in color than the remainder of the coated blade and contain an excess or “bubbled” material inside the round spot. These “white spots” appear to form upon application of the seal coat. Coated blades using the prior art multilayer overlay system may also exhibit a “picture frame” effect with the layers being thicker near the blade edges, thus leading to weaker overlay adhesion and likely edge peeling. All these defects being irregularities in the sealed overlay surface not only reduce aerodynamic efficiency of the blade, but also might serve as active sites for thermal and corrosion attack.
In view of the above-identified concerns and disadvantages, a need exists for continuous improvements to the surface finish characteristics as well as thermal and corrosive performance of the prior art slurry-based, multilayer overlay systems. While the prior art slurry-based, multilayer overlay systems meet the requirements and specifications of current engine manufacturers, improvements are needed for use with newer, more advanced engines. It would therefore be desirable to provide a multilayer overlay system that improves upon the surface finish characteristics of the prior art overlay systems and possesses improved thermal stability in normal and corrosive environments.